The fuselages of many aircraft consist of circumferential frame members, longitudinal stringers, and a thin skin, all made from lightweight aluminium. This construction allows for a balance of flight properties and weight.
The sheets of aluminium that make up the skin are connected together as lap joints by generally two to three rows of rivets. The outer skin later is countersunk at each rivet location so the rivet head is flush with the skin, resulting in optimal aerodynamic properties.
When the skin is subjected to the stresses of normal operation, particularly in pressurized commercial aircraft, fatigue damage can occur in the metal sheets and especially in high stress locations around fasteners. The problem is exacerbated by the ingress of environmental elements and leads to the joint cracking. Crack growth, if left undetected, can lead to catastrophic failure, as in the case of Aloha Airlines Flight 243 in 1988. As the aircraft reached its normal flight altitude of 24,000 feet (7,300 m), a small section on the left side of the roof ruptured. The resulting explosive decompression tore off a large section of the roof, consisting of the entire top half of the aircraft skin extending from just behind the cockpit to the fore-wing area. It was subsequently discovered that the incident was caused by the presence of multiple small cracks which arose as a result of environmental degradation of the joint located aft of the front port side passenger door. This phenomenon has subsequently been termed “multi-site damage” (MSD).
Since the Aloha incident, aircraft operators have been directed to regularly check for the presence of cracks and MSD in the fuselage skin. In order to identify the presence of cracks before they reach critical lengths, various inspection techniques are utilized.
While visual inspection is an important part of the detection process, however many naturally occurring cracks in their initiation are simply too small to see or otherwise detect. To assist with the detection of these small and hidden cracks, non-destructive inspection (NDI) methods are used. NDI methods can also be used to detect cracks that exist under paint and detect areas of corrosion between the layers of skin. Some of the more common NDI methods used in aircraft fuselage crack detection are ultrasound and eddy current methods. These methods are not capable of detecting all cracks and are particularly poor in detecting small naturally occurring defects.
After the Aloha Airlines Flight 243 accident, all 737's with over 50,000 cycles we required to have their lap joints reinforced with external sheet metal patches. This modification is costly, and takes about 15,000 man hours.
The Alohoa accident further highlighted the problem of multiple interacting repairs. Despite the presence of MSD in the fuselage lap joint the failure in that incident ran from corrosion repair to corrosion repair. Indeed, for Boeing 727 aircraft [42, 43] there are numerous instances where no crack growth was noted until after a corrosion repair had been installed.
The problem of multiple interacting repairs to corrosion damage is not confined to corrosion in fuselage lap joints. The common approach to corrosion damage in operational aircraft is to cut out the corrosion and rivet a mechanical doubler over the region. Unfortunately if the aircraft is operated in an aggressive environment then corrosion is likely to occur over a (relatively) broad area and this can lead to a number of mechanical repairs that lie in relatively close proximity. This repair process involves drilling holes, which act as stress concentrators, in the base structure and unless the operational environment changes these holes provide sites at which corrosion pits can develop and subsequently crack as was seen in the Aloha accident.
As such a repair methodology is needed whereby the structure need not be further damaged and new sites at which pitting and subsequent cracking are created.
As another example, in April 2011 a fuselage lap joint in a Southwest Airlines Boeing 737-300 aircraft tore an 18 inch hole in the roof, and led to the grounding of 79 of its older Boeing 737 aircraft for inspections [38, 39]. This resulted in the cancelation of almost 700 flights [38, 39]. These inspections, which found cracks in a total of four Southwest aircraft, [38] led to the US FAA mandating the inspection of 175 737 aircraft that had experienced more than 35,000 cycles. There are more than 931 similar aircraft worldwide. The problem is not confined to 737 and 727 aircraft. On 26th October 2010 an American Airlines 757-200 aircraft was forced to land at Miami International Airport due to a sudden decompression arising from cracking in a fuselage joint. This aircraft had experienced less than 23,000 cycles. This led to the discovery of cracking in other 757 aircraft and a January 2011 FAA Airworthiness directive [40] mandating the inspection of all 757-200 and 757-300 aircraft.
Environmental degradation and subsequent crack initiation and progression is not just of importance to commercial airlines. Military aircraft, particularly those with advanced age, can also develop environmental degradation and cracking at fastener holes. As the military attempts to keep its fleet flight-worthy for longer periods of time, additional prevention, inspection, and mitigation methods are being developed to prevent both environmental degradation and catastrophic failure.
When cracks are discovered, they are typically repaired by the application of external sheet metal patches. Again, this is a costly and time consuming process. A further problem is that the application of patches may actually initiate a weakness in the underlying structure. Such undetected and undetectable cracks can compromise the safety of the fuselage/wing skin. These repairs can also locally over-stiffen the structure and result in catastrophic failure in the fuselage/wing skin as a result of a crack running from repair to repair.
Externally bonded composite bonded repairs have been developed to address this problem. However, these repairs do not prevent the ingress of moisture and hence do not alleviate environmental degradation of the structure. Furthermore, to ensure a durable bond the structure needs to be heated to approximately 120 C. Additionally, composite bonded repairs cannot be used in regions where there is a tight radius of curvature.
It is an aspect of the present invention to overcome or alleviate a problem of the prior art by providing a method for preventing or repairing a structural weakness in an aircraft structure. A further aspect of the present invention is to overcome or alleviate a problem of the prior art by providing a method for preventing environmental degradation in an aircraft structure.
The discussion of documents, acts, materials, devices, articles and the like is included in this specification solely for the purpose of providing a context for the present invention. It is not suggested or represented that any or all of these matters formed part of the prior art base or were common general knowledge in the field relevant to the present invention as it existed before the priority date of each provisional claim of this application.